Combustor liner panel end rail with diffused interface passage for a gas turbine engine combustor

ABSTRACT

A liner panel for use in a combustor of a gas turbine engine, the liner panel including a cold side; and a rail that extends from the cold side, the rail includes a first diffusion interface passage surface and a second diffusion interface passage surface, the first diffusion interface passage surface angled with respect to the second diffusion interface passage surface.

BACKGROUND

The present disclosure relates to a gas turbine engine combustor and,more particularly, to a liner panel therefor.

Gas turbine engines, such as those that power modern commercial andmilitary aircraft, generally include a compressor section to pressurizean airflow, a combustor section to burn a hydrocarbon fuel in thepresence of the pressurized air, and a turbine section to extract energyfrom the resultant combustion gases.

Among the engine components, relatively high temperatures are observedin the combustor section such that cooling airflow is provided to meetdesired service life requirements. The combustor section typicallyincludes an annular combustion chamber formed by an inner and outer wallassembly. Each wall assembly includes a support shell lined with heatshields, which are often referred to as liner panels. The liner panelsmay be segmented to accommodate thermal growth in operation and forother considerations. The combustor liner panels include a hot sideexposed to the gas path. The opposite, or cold side, has features suchas threaded studs to mount the liner panel to the support shell, and aperimeter rail that contacts the inner surface of the support shells.

The liner panel perimeter rail includes a forward rail that forms theforward, or upstream, edge of the panel, an aft rail that forms the aft,or downstream, edge of the liner panel, and longitudinal side rails thatconnect the forward and aft rails.

The liner panels extend over an arc in a conical or cylindrical arrayand axially interface in regions where the combustor geometrytransitions (e.g., diverges or converges). This interface may contributeto durability and flow path concerns where forward and aft, as well ascircumferentially adjacent, panels abut. These interfaces may be proneto steps between panels, dead regions, cooling challenges and adverselocal aerodynamics.

SUMMARY

A liner panel for use in a combustor of a gas turbine engine, the linerpanel according to one disclosed non-limiting embodiment of the presentdisclosure includes a rail that extends from a cold side, the railincludes a first diffusion interface passage surface and a seconddiffusion interface passage surface, the first diffusion interfacepassage surface angled with respect to the second diffusion interfacepassage surface.

A further aspect of the present disclosure includes, wherein the firstdiffusion interface passage surface at least partially defines apre-diffuser section and the second diffusion interface passage surfaceat least partially defines a diffuser section along one side of adiffusion interface passage axis.

A further aspect of the present disclosure includes, wherein thediffuser section of the rail defines a height with respect to the coldside between 0.05 H-0.6 H, where H is the height of the rail.

A further aspect of the present disclosure includes, wherein thediffuser section at least partially defines an expanding passage thatextends at an angle between 1-15 degrees with respect to a surface thatforms the pre-diffuser section.

A further aspect of the present disclosure includes, wherein the linerpanel is at least one of a forward liner panel, and an aft liner panel.

A further aspect of the present disclosure includes, wherein the rail isa periphery rail that defines an edge of the liner panel.

A further aspect of the present disclosure includes, wherein the railincludes a ramped end section.

A further aspect of the present disclosure includes, wherein the rail istrapezoidal shaped.

A further aspect of the present disclosure includes, wherein the rail isramp shaped.

A combustor for a gas turbine engine according to one disclosednon-limiting embodiment of the present disclosure includes a supportshell; a first liner panel mounted to the support shell, the first linerpanel including a first rail that extends from a cold side of the firstliner panel, the first rail includes a diffusion interface passagesurface that is non-perpendicular to the hot side; and a second linerpanel mounted to the support shell, the second liner panel including asecond rail that extends from a cold side of the second liner panel, thesecond rail adjacent to the first rail to at least partially form adiffusion interface passage.

A further aspect of the present disclosure includes a multiple of studsthat extend from the first liner panel, the first liner panel mounted tothe support shell via the multiple of studs, and a multiple of studsthat extend from the second liner panel, the second liner panel mountedto the support shell via the multiple of studs.

A further aspect of the present disclosure includes, wherein the firstliner panel is a forward liner panel.

A further aspect of the present disclosure includes, wherein the secondliner panel is an aft liner panel.

A further aspect of the present disclosure includes, wherein thediffusion interface passage defines a diffusion interface passage axisoriented at 30-90 degrees with respect to a hot side of the second linerpanel.

A further aspect of the present disclosure includes, wherein thediffusion interface passage defines a pre-diffuser section and adiffuser section along the diffusion interface passage axis.

A further aspect of the present disclosure includes, wherein thediffuser section of the first rail defines a height with respect to acold side of the first liner panel between 0.05-H 0.6 H, where H is theheight of the first rail.

A further aspect of the present disclosure includes, wherein thediffuser section of the first rail defines an expanding passage orientedat an angle between 1-15 degrees from an aft diffusion interface passagesurface of the first rail.

A further aspect of the present disclosure includes, wherein thediffuser section of the second rail defines a height with respect to acold side of the second liner panel between 0.05 H-0.6 H, where H is theheight of the second rail.

A further aspect of the present disclosure includes, wherein thediffuser section of the second rail defines an expanding passageoriented at an angle between 1-15 degrees from a forward diffusioninterface passage surface of the second rail.

A further aspect of the present disclosure includes, wherein the firstrail is trapezoidal shaped.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of an example gas turbine enginearchitecture;

FIG. 2 is an expanded longitudinal schematic sectional view of acombustor section according to one non-limiting embodiment that may beused with the example gas turbine engine architectures;

FIG. 3 is an exploded partial sectional view of a portion of a combustorwall assembly;

FIG. 4 is a perspective cold side view of a portion of an outer linerpanel array;

FIG. 5 is a perspective partial sectional view of the combustor;

FIG. 6 is a sectional view of a portion of a combustor wall assembly;

FIG. 7 is a sectional view of a combustor wall assembly with a diffusioninterface passage according to one embodiment;

FIG. 8 is a sectional view of a combustor wall assembly with a diffusioninterface passage according to another embodiment; and

FIG. 9 is a sectional view of a combustor wall assembly with a diffusioninterface passage according to still another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative enginearchitectures might include an augmenter section among other systems orfeatures. The fan section 22 drives air along a bypass flowpath and intothe compressor section 24. The compressor section 24 drives air along acore flowpath for compression and communication into the combustorsection 26, which then expands and directs the air through the turbinesection 28. Although depicted as a turbofan in the disclosednon-limiting embodiment, it should be appreciated that the conceptsdescribed herein are not limited to use with turbofans as the teachingsmay be applied to other types of turbine engines such as a turbojets,turboshafts, and three-spool (plus fan) turbofans.

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearing systems38. The low spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor (“LPC”) 44, and a lowpressure turbine (“LPT”) 46. The inner shaft 40 drives the fan 42directly or through a geared architecture 48 to drive the fan 42 at alower speed than the low spool 30. An exemplary reduction transmissionis an epicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor (“HPC”) 52 and a high pressure turbine (“HPT”) 54. Acombustor 56 is arranged between the HPC 52 and the HPT 54. The innershaft 40 and the outer shaft 50 are concentric and rotate about theengine central longitudinal axis A which is collinear with theirlongitudinal axes.

Core airflow is compressed by the LPC 44, then the HPC 52, mixed withthe fuel and burned in the combustor 56, then expanded over the HPT 54and the LPT 46. The LPT 46 and HPT 54 rotationally drive the respectivelow spool 30 and high spool 32 in response to the expansion.

In one non-limiting example, the geared architecture 48 has a gearreduction ratio of greater than about 2.3, and, in another example, isgreater than about 2.5:1. The geared turbofan enables operation of thelow spool 30 at higher speeds which can increase the operationalefficiency of the LPC 44 and LPT 46 to provide increased pressure in afewer number of stages. A pressure ratio associated with the LPT 46 ispressure measured prior to the inlet of the LPT 46 as related to thepressure at the outlet of the LPT 46 prior to an exhaust nozzle of thegas turbine engine 20. In this example, the bypass ratio of the gasturbine engine 20 is greater than about ten (10:1), the fan diameter issignificantly larger than that of the LPC 44, and the LPT 46 has apressure ratio that is greater than about five (5:1).

A significant amount of thrust is provided by the bypass flow due to thehigh bypass ratio. The fan section 22 of the gas turbine engine 20 isdesigned for a particular flight condition—typically cruise at about 0.8Mach and about 35,000 feet (10668 m). This flight condition, with thegas turbine engine 20 at its best fuel consumption, is also referred toas bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is anindustry standard parameter of fuel consumption per unit of thrust. Fanpressure ratio is the pressure ratio across a blade of the fan section22 without the use of a fan exit guide vane (FEGV) system. The low fanpressure ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low corrected fan tip speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (“Tram”/518.7)^(0.5) and according to this example is lessthan about 1150 fps (351 m/s). It should be appreciated that the aboveparameters are only exemplary of one geared architecture engine and thatthe present disclosure is applicable to other gas turbine enginesincluding direct drive turbofans.

With reference to FIG. 2, the combustor section 26 generally includes acombustor 56 with an outer combustor wall assembly 60, an innercombustor wall assembly 62, and a diffuser case module 64. The outercombustor wall assembly 60 and the inner combustor wall assembly 62 arespaced such that an annular combustion chamber 66 is definedtherebetween to surround the engine central longitudinal axis A.

The outer combustor wall assembly 60 is spaced radially inward from anouter diffuser case 64A of the diffuser case module 64 to define anouter annular plenum 76. The inner combustor wall assembly 62 is spacedradially outward from an inner diffuser case 64B of the diffuser casemodule 64 to define an inner annular plenum 78. Although a particularcombustor is illustrated, other combustor types with various combustorliner arrangements will also benefit herefrom.

The combustor wall assemblies 60, 62 contain the combustion products fordirection toward the turbine section 28. Each combustor wall assembly60, 62 includes a respective outer support shell 68 and inner supportshell 70 which support one or more liner panels 72, 74 mounted theretoand arranged to form a liner panel array. The support shells 68, 70 maybe manufactured by, for example, the hydroforming of a sheet metal alloyto provide the generally cylindrical outer shell 68 and inner shell 70.Each of the liner panels 72, 74 may have a quadrilateral projection(e.g., rectangular or trapezoidal). The liner panels 72, 74 may bemanufactured of, for example, a nickel based super alloy, ceramic, orother temperature resistant material. In one disclosed non-limitingembodiment, the liner panel array includes a multiple of forward outerliner panels 72A and a multiple of aft outer liner panels 74A that arecircumferentially arranged to line the outer shell 68. A multiple offorward inner liner panels 72B and a multiple of aft inner liner panels74B are circumferentially arranged to line the inner shell 70.

The combustor 56 further includes a forward assembly 80 downstream ofthe compressor section 24 to receive compressed airflow therefrom. Theforward assembly 80 generally includes a cowl 82, a bulkhead assembly84, and a multiple of swirlers 90. Each of the swirlers 90 is alignedwith one of a multiple of fuel nozzles 86 and a respective hood port 94along an axis F.

The cowl 82 extends between, and is secured to, a forward section of thecombustor wall assemblies 60, 62. The cowl 82 includes a multiple ofhood ports 94 that each receive one of the respective multiple of fuelnozzles 86 and facilitates the direction of compressed air into theforward section of the combustion chamber 66 through a swirler opening92. Each fuel nozzle 86 may be mounted to the diffuser case module 64 toproject through one hood port 94 and respective swirler 90 mounted inthe bulkhead assembly 84.

The bulkhead assembly 84 includes a bulkhead support shell 96 securedtransversely to the combustor wall assemblies 60, 62, and a multiple ofbulkhead liner panels 98 secured to the bulkhead support shell 96 tosurround each swirler 90. The bulkhead support shell 96 is formed of amultiple of bulkhead liner panels 98 that are segmented, typically oneto each fuel nozzle 86 and swirler 90.

The forward assembly 80 introduces a portion of the compressed air fromthe HPC 24 into the forward section of the combustion chamber 66 whilethe remainder enters the outer annular plenum 76 and the inner annularplenum 78. The multiple of fuel nozzles 86 and swirlers 90 produce ablended fuel-air mixture that supports stable combustion in thecombustion chamber 66.

Opposite the forward assembly 80, the combustor wall assemblies 60, 62terminate adjacent to a first row of Nozzle Guide Vanes (NGVs) 54A inthe HPT 54. The NGVs 54A are static engine components which direct coreairflow combustion gases onto turbine blades of the first turbine rotorin the turbine section 28 to facilitate the conversion of core airflowcombustion gas pressure energy into kinetic energy. The core airflowcombustion gases are also accelerated by the NGVs 54A, and due to theconvergent shape, facilitate a “spin” or a “swirl” in the direction ofHPT 54 rotation. The turbine rotor blades receive this energy to drivethe turbine rotor at high speed.

With reference to FIG. 3, a multiple of studs 100 extend from each ofthe liner panels 72, 74 so as to permit a liner array (partially shownin FIG. 4) of the liner panels 72, 74 to be mounted to their respectivesupport shells 68, 70 with fasteners 102 such as nuts. That is, thestuds 100 project rigidly from a cold side 110 of the liner panels 72,74 to extend through the respective support shells 68, 70 and receivethe fasteners 102 on a threaded section of the studs 100 (FIG. 5).

A multiple of cooling impingement passages 104 penetrate through thesupport shells 68, 70 to allow air from the respective annular plenums76, 78 to enter wall cavities 106 formed in the combustor wallassemblies 60, 62 between the respective support shells 68, 70 and linerpanels 72, 74. The impingement passages 104 are generally normal to thesurface of the liner panels 72, 74. The air in the wall cavities 106provides cold side impingement cooling of the liner panels 72, 74 thatis generally defined herein as heat removal via internal convection.

A multiple of effusion passages 108 penetrate through each of the linerpanels 72, 74. The geometry of the passages, e.g., diameter, shape,density, surface angle, incidence angle, etc., as well as the locationof the passages with respect to the high temperature combustion flowalso contributes to effusion cooling. The effusion passages 108communicate air from the wall cavities 106 defined in part by the coldside 110 of the liner panels 72, 74 to a hot side 112 of the linerpanels 72, 74, and thereby facilitate the formation of a thin,relatively cool, film of cooling air along the hot side 112. Theeffusion passages 108 are generally more numerous than the impingementpassages 104 and promote film cooling along the hot side 112 to sheaththe liner panels 72, 74 (FIG. 6). Film cooling as defined herein is theintroduction of a relatively cooler air at one or more discretelocations along a surface exposed to a high temperature environment toprotect that surface in the region of the air injection as well asdownstream thereof. It should be further appreciated that the disclosedcooling flow paths as shown schematically by arrows are but anillustrated example and should not be limited only thereto.

A multiple of dilution passages 116 are located thorough the linerpanels 72, 74 and the respective support shells 68, 70, each along acommon axis D. For example, the dilution passages 116 are located alonga circumferential line (illustrated schematically by line W; shownpartially in FIG. 4). Although the dilution passages 116 are illustratedin the disclosed non-limiting embodiment as within the aft liner panels74A, 74B, the dilution passages may alternatively be located in theforward liner panels 72A, 72B or in a single liner panel which replacesthe fore/aft liner panel array.

With reference to FIG. 4, in one disclosed non-limiting embodiment, eachof the forward liner panels 72A, 72B, and the aft liner panels 74A, 74Bin the liner panel array includes a respective perimeter rail 120A,120B, each respectively formed by a forward circumferential rail 122 a,122 b, an aft circumferential rail 124 a, 124 b, and axial rails 126Aa126Ab, 126Ba, 126Bb, that interconnect the respective forward and aftcircumferential rail 122 a, 122 b, 124 a, 124 b. The perimeter rails120A, 120B seal each liner panel with respect to the respective supportshell 68, 70 to form the impingement cavity 106 therebetween. That is,the forward and aft circumferential rails 122 a, 122 b, 124 a, 124 b arelocated at relatively constant curvature shell interfaces while theaxial rails 126Aa, 126Ab, 126Ba, 126Bb, extend along an axial length ofthe respective support shell 68, 70 to complete the perimeter rail 120A,120B to seal the forward liner panel 72A, 72B and the aft liner panel74A, 74B to the respective support shell 68, 70. Although only perimeterrails 120A, 120B of respective forward and aft liner panels 72A, 74A aredescribed in detail, it should be appreciated that each liner paneltypically includes such a perimeter rail.

The multiple of studs 100 are located adjacent to the respective forwardand aft circumferential rail 122 a, 122 b, 124 a, 124 b. Each of thestuds 100 may be at least partially surrounded by posts 130 to supportthe fastener 102 and provide a stand-off between the forward linerpanels 72A, 72B, the aft liner panels 74A, 74B, and respective supportshell 68, 70. The dilution passages 116 are located downstream of theforward circumferential rail 122 a, 122 b in the aft liner panels 74A,74B to quench the hot combustion gases within the combustion chamber 66by direct supply of cooling air from the respective annular plenums 76,78. That is, the dilution passages 116 pass air at the pressure outsidethe combustion chamber 66 directly into the combustion chamber 66. Thisdilution air is not primarily used for cooling of the metal surfaces ofthe combustor shells or panels, but to condition the combustion productswithin the combustion chamber 66.

With reference to FIG. 7, in one embodiment, the aft circumferentialrail 124 a of the forward liner panel 72A is adjacent to the forwardcircumferential rail 122 b of the aft liner panel 74A to form adiffusion interface passage 150 oriented at an angle to the gas pathflow (illustrated schematically by arrow G) through the combustor 66.Although only the radially outer forward liner panel 72A and theradially outer aft liner panel 74A are illustrated in this example, toform the circumferentially arranged diffusion interface passage 150, thediffusion interface passage 150 may alternatively, or additionally, belocated between the radially inner forward liner panel 72B and aft linerpanel 74B, and/or axially between the respective radially outer forwardouter liner panels 72A, radially outer aft liner panel 74A, radiallyinner forward liner panels 72B, and/or radially inner aft liner panel74B.

The diffusion interface passage 150 forms a slot interface between therespective forward outer liner panel 72A and the aft outer liner panel74A. The aft circumferential rail 124 a and the forward circumferentialrail 122 b may be machined, forged, casted, additively manufactured, orotherwise formed.

In this embodiment, the diffusion interface passage 150 is oriented atan angle between 30-90 degrees with respect to the hot side 112 of theradially outer aft liner panel 74A. The diffusion interface passage 150includes a pre-diffuser section 152 and a diffuser section 154. Thediffusion interface passage 150 is angled in a direction to ejectairflow therethrough generally with the gas path flow G through thecombustor 66.

The aft circumferential rail 124 a of the forward liner panel 72Aincludes a first aft diffusion interface passage surface 156 of thepre-diffuser section 152, and a second aft diffusion interface passagesurface 158 of the diffuser section 154, the first aft diffusioninterface passage surface 156 angled relative to the second aftdiffusion interface passage surface 158. The forward circumferentialrail 122 b of the aft liner panel 74A also includes a first forwarddiffusion interface passage surface 160 of the pre-diffuser section 152and a second forward diffusion interface passage surface 162 of thediffuser section 154, the first forward diffusion interface passagesurface 160 angled relative to the second forward diffusion interfacepassage surface 162.

The aft circumferential rail 124 a of the respective forward outer linerpanel 72A and the forward circumferential rail 122 b of the aft linerpanel 74A extends for a height H with respect to the cold side 110. Inone example, the diffuser section 154 defines a height with respect tothe cold side 110 between 0.05 H-0.6 H.

The diffuser section 154 is defined as an expanding passage that expandsat an angle between 1-15 degrees between the first aft diffusioninterface passage surface 156 and the second aft diffusion interfacepassage surface 158 of the aft circumferential rail 124 a, and 1-15degrees from the first forward diffusion interface passage surface 160and the second forward diffusion interface passage surface 162 of theforward circumferential rail 122 b.

With reference to FIG. 8, in another embodiment, the aft circumferentialrail 124 a includes a forward surface 160 to form a trapezoidal shapedrail 124 a in cross-section. Other shapes that facilitatemanufacturability may alternatively or additionally be provided.

With reference to FIG. 9, in another embodiment, the aft circumferentialrail 124 a is ramp shaped in cross-section. The forward circumferentialrail 122 b of the aft liner panel 74A, 74B includes a ramped end section170 but is otherwise of the same thickness as the liner panel 74Aitself. That is, the ramped end section 170 is chiseled or otherwisecontoured to provide the desired diffusion interface passage 150.

The diffusion interface passage 150 ejects airflow generally along thehot side 112 of the liner panel 74 to promote film attachment to the hotside 112 and facilitate cooling effectiveness. The diffusion interfacepassage 150 also shields the support shell 68, 70 from a line of sightheat flow path to the combustion chamber 66. This facilitates combustordurability and time on wing.

The use of the terms “a” and “an” and “the” and similar references inthe context of description (especially in the context of the followingclaims) are to be construed to cover both the singular and the plural,unless otherwise indicated herein or specifically contradicted bycontext. The modifier “about” used in connection with a quantity isinclusive of the stated value and has the meaning dictated by thecontext (e.g., it includes the degree of error associated withmeasurement of the particular quantity). All ranges disclosed herein areinclusive of the endpoints, and the endpoints are independentlycombinable with each other. It should be appreciated that relativepositional terms such as “forward,” “aft,” “upper,” “lower,” “above,”“below,” and the like are with reference to the normal operationalattitude of the vehicle and should not be considered otherwise limiting.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beappreciated that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A liner panel for use in a combustor of a gasturbine engine, the liner panel comprising: a cold side; and a rail thatextends from the cold side, the rail includes a first diffusioninterface passage surface and a second diffusion interface passagesurface, the first diffusion interface passage surface angled withrespect to the second diffusion interface passage surface.
 2. The linerpanel as recited in claim 1, wherein the first diffusion interfacepassage surface at least partially defines a pre-diffuser section andthe second diffusion interface passage surface at least partiallydefines a diffuser section along one side of a diffusion interfacepassage axis.
 3. The liner panel as recited in claim 2, wherein thediffuser section of the rail defines a height with respect to the coldside between 0.05 H-0.6 H, where H is the height of the rail.
 4. Theliner panel as recited in claim 3, wherein the diffuser section at leastpartially defines an expanding passage that extends at an angle between1-15 degrees with respect to a surface that forms the pre-diffusersection.
 5. The liner panel as recited in claim 1, wherein the linerpanel is at least one of a forward liner panel, and an aft liner panel.6. The liner panel as recited in claim 1, wherein the rail is aperiphery rail that defines an edge of the liner panel.
 7. The linerpanel as recited in claim 1, wherein the rail includes a ramped endsection.
 8. The liner panel as recited in claim 1, wherein the rail istrapezoidal shaped.
 9. The liner panel as recited in claim 1, whereinthe rail is ramp shaped.
 10. A combustor for a gas turbine enginecomprising: a support shell; a first liner panel mounted to the supportshell, the first liner panel including a first rail that extends from acold side of the first liner panel, the first rail includes a diffusioninterface passage surface that is non-perpendicular to the hot side; anda second liner panel mounted to the support shell, the second linerpanel including a second rail that extends from a cold side of thesecond liner panel, the second rail adjacent to the first rail to atleast partially form a diffusion interface passage.
 11. The combustor asrecited in claim 10, further comprising a multiple of studs that extendfrom the first liner panel, the first liner panel mounted to the supportshell via the multiple of studs, and a multiple of studs that extendfrom the second liner panel, the second liner panel mounted to thesupport shell via the multiple of studs.
 12. The combustor as recited inclaim 10, wherein the first liner panel is a forward liner panel. 13.The combustor as recited in claim 10, wherein the second liner panel isan aft liner panel.
 14. The combustor as recited in claim 10, whereinthe diffusion interface passage defines a diffusion interface passageaxis oriented at 30-90 degrees with respect to a hot side of the secondliner panel.
 15. The combustor as recited in claim 14, wherein thediffusion interface passage defines a pre-diffuser section and adiffuser section along the diffusion interface passage axis.
 16. Thecombustor as recited in claim 15, wherein the diffuser section of thefirst rail defines a height with respect to a cold side of the firstliner panel between 0.05 H-0.6 H, where H is the height of the firstrail.
 17. The combustor as recited in claim 16, wherein the diffusersection of the first rail defines an expanding passage oriented at anangle between 1-15 degrees from an aft diffusion interface passagesurface of the first rail.
 18. The combustor as recited in claim 15,wherein the diffuser section of the second rail defines a height withrespect to a cold side of the second liner panel between 0.05 H-0.6 H,where H is the height of the second rail.
 19. The combustor as recitedin claim 18, wherein the diffuser section of the second rail defines anexpanding passage oriented at an angle between 1-15 degrees from aforward diffusion interface passage surface of the second rail.
 20. Thecombustor as recited in claim 15, wherein the first rail is trapezoidalshaped.